Missile directional control system

ABSTRACT

1. An attitude control system for maintaining a missile on a launch predetermined path comprising means mounted on the missile for indicating a deviation from said predetermined path including means for providing an output signal representative of angular position change and means for providing an output signal representative of angular rate change, means for modifying said angular position change signal by multiplying it by a factor &#39;&#39;&#39;&#39;a&#39;&#39;&#39;&#39; whereby a congruent ((T-D)/(k/p) - 1), an algebraic summation means responsive to said modified position and rate change signals and operable to provide a combined output representative of their algebraic summation, and means operatively connected to and controlled by the output of said summation means, said last mentioned means operable to provide a corrective force to return the missile to its predetermined path.

United States Patent H June 24, 1975 May 1 MISSILE DIRECTIONAL CONTROL SYSTEM [75] Inventor: Robert F. May, Birmingham. Mich. [731 Assignee: Chrysler Corporation, Highland Park, Mich.

[221 Filed: May 28, 1963 [21] Appl. No.: 286.100

(52] US. Cl. 244/3.22 [51] Int. Cl. F42b 15/02 [58] Field 01' Search 244/14, 14.1-14.6,

[56] References Cited UNITED STATES PATENTS 2,873,4l3 2/1959 Owcn 244/77 D 2.974.594 3/1961 Bochm 102/50 Primary !.\'umincrVerlin R. Pendegrass Almrm-y. Agent, or FirmTalburtt & Baldwin EXEMPLARY CLAIM 1. An attitude control system for maintaining a missile on a launch predetermined path comprising means mounted on the missile for indicating a deviation from said predetermined path including means for providing an output signal representative of angular position change and means for providing an output signal representative of angular rate change, means for modifying said angular position change signal by multiplying it by a factor *a" whereby a '--l(T-D)l(k/p) 4 1}, an algebraic summation means responsive to said modilied position and rate change signals and operable to provide a combined output representative of their a1 gebraic summation, and means operatively connected to and controlled by the output of said summation means. said last mentioned means operable to provide a corrective force to return the missile to its predetermined path.

16 Claims, 9 Drawing Figures fat 1,7

SHEET PATENTEDJUN 24 I975 INVENTOR.

%Z67'2 I W WW DW PATENTED JUN 2 4 I975 SHEET INVENTOR. F0267; BY

HTTdW/VJ/ MISSILE DlRl'lUl'lONAl, (ONTROL SYS'I'EM This invention relates generally to missile directional control systems and. more particularly. to a directional control system to provide missile guidance during its high acceleration boost flight.

ln provision of a directional control system. a primary requirement is that the missile be aligned in the correct direction of travel at the end of boost llight. ()thcrwisc stated. there should be no velocity or displacement pcrpmidicular to the direction in which the missile was initially aimed and launched. The boost flight phase may then be followed by free flight. subject to ballistic wind effects. or it can be followed by a coir trolled sustained flight. When the directional control system is utilized during the entire missile llight. the primary requirement is that the lateral displacement at the end of flight he held to an absolute minimum.

It is an obiect of the present invention to provide a directional control system operable without external control data and utilizing body-fixed attitude angle and angular rate information to exercise its control function.

It is an additional object of this invention to provide a directional control system and control method which requires no missile spin for stability of operation but is operable in the presence of such spin.

It is a further object of this invention to provide a directional control system providing attitude corrective thrust forces of substantially linear variation with esceedingly short time lag behind the initiating error signal.

The above objects are best realized by a directional control system which measures the pitch and yaw of a missile which is aerodynamically stable through the Mach number range under which it operates under directional control and then applies torques to restrain this motion in such a manner that motor thrust forces. aerodynamic forces and control forces remain balanced in directions perpendicular to the flight path. The present invention. first briefly described. comprises a simple two-degrce-offreedom gyroscope. capable of operation under high acceleration and mounted with its spin axis along the missile longitudinal axis. Electrical outputs are taken from gyro pick-offs which are representative of angular deviation and rate of deviation ofthe missile body with respect to the spin axis. 'l'hcse outputs are utilized in an electronic com puter which linearly combines this data and produces an electrical control output. (ontrol jets are fixed to the missile in the same body plane as the gyro pick-offs and are adapted to provide corrective forces normal to the missile body. The electrical control output is employed to direct corrective forces through these jets into the airstrcam or. alternately. into the boost rocket ltU/JlL.

l'o facilitate the understanding of the present invention. it may be characterized as being of the open-loop type i.e. one in which no direct measurement is made of the lateral velocity or displacement of the missile flight path which is to he minimived. The control gains are functions of aerodynamic characteristics. thrust level. and missile geometry. Hy straightforward static analysis. it will be seen that the magnitude of the cow trol force is proportional to the quantity (thrust minus longitudinal drag) times the pitch or yaw angle measured by the gyro. divided by the quantity ratio ofcontrol moment arm to aerodynamic moment arml minus (it l l and the control equation maybe expressed as follows:

in which .I control force 'l thrust l) drag k control moment arm P aerodynamic moment arm (I): pitch or yaw angle a gain coefficient Under conditions of boost acceleration and high thrust the i -D term is relatively constant. As the missile propellant is used up. k varies with the changing missile weight. The quantity p varies with the Mach number.

It will be observed that the above control equation is undamped. It is desirable to incorporate a rate term to provide a damping function to render the control force a function of both the angular term and an angular rate term as follows:

J u d) +l1ib' in which I) if) is the angular rate term. The angular rate term I if) is a function of missile inertia and control transfer function.

In the presence of a missile roll. a cross coupling exists between the pitch and yaw controls due to a roll effeet on angular rate. If the missile roll is significant. the control equation may be modified to include a third term as follows:

in which gain coefficient 0 roll angular rate 6 gyro angle in the other channel It should be noted that the term 6 refers to pitch angle for the yaw control channel or to negative yaw angle for the pitch control channel. It is generally desirable to utilize constant values for a. b. and 1' rather than to program each. although employment of a program may be made to increase flight path accuracy.

in view of the foregoing objects and comments. further understanding of the present invention will become apparent from a consideration of the accompanying specifications and drawings in which Flti. l is a functional block diagram of the directional control system incorporating the present invention;

FIG. 2 is a schematic drawing of the electronic computer which performs the operations defined by the control equations previously noted;

Fl(i. 3 is a cross sectional view of the servo regulator valve utilized:

Flti. 4 is a bottom plan view of the manifold block utilized in combination with the servo valve of Fl(i. 3-.

FIG. 5 is a sectional view taken along the line AA of FIG. 4 and includes a diagrammatic showing of the servo valve in its operative relationship to the manifold block.

l-'l(i. 6 is a sectional view taken along the line B-B of Flti. 4'. and

FIG. 7 is a sectional view taken along the line (-C of HG. 4;

FIG. 8 is a schematic drawing of position pickoff l2. and

FIG. 9 is a schematic drawing view of a rate plLltttff l4.

In FIG. I. a control system fora missile is shown with reference to one plane of control. as for example. in the pitch plane. It will be noted that in the practice of the present invention an additional system would be employed in the yaw control plane also. which would be essentially a duplicate of the system hereinafter shown and described. lncluded as an element of this system is a gyroscope ll) which may be a two degrec-of-freedom gyroscope Capable of operation under high accelera' tion and mounted with its spin axis along the missile longitudinal axis. Gyroscope may be of the spherical air bearing type with the inertial element. namely. its rotor. external to the bearing. An angular position pick off 12 and an angular rate pickoff M are located at spaced points on the periphery of the rotating inertial element and are operable to provide electrical outputs representative of the magnitude ofangular position and angular rate changes. respectively. The outputs from pickoffs l2 and 14 are fed into an electronic computer 16. More specifically. the output from position pieltoff l2 may be directed through a pre-amplifier [8 into a summing amplifier 20. The output from rate pickoff 14 is directed through a modulator 2] and preamplifier 22 to summing amplifier stage 20. The summed signal outputs from summing amplifier pass through a demodulator 24. A DC. power source 28 is utilized to furnish power for the several aforementioned amplificrs and stages utilized in the system. A static inverter 30 is utilized to provide a constant reference signal to demodulator 24 and for excitation of the position pickoffs. The amplified d.c. outputs are then directed in the form of pulses of plus or minus polarity to the control coils 32 of servo valve 34. Depending on whether the control signal output of the computer 16 is plus or minus, one of the oppositely oriented thrust jets 36.38 will be operated with a corrective force F of the appropri ate magnitude. Thrust jets 36.38 are controlled in their action by servo valve 34. a manifold block operatively associated therewith, and servo regulators 42 and 44, respectively. The operating pressure for servo valve 34 and servo regulators 42,44 are derived from a common source of pressurized fluid 46. Inserted in the line between servo valve 34 and pressure source 46 is a regulator valve 48 of a type well known in the art whose function is to maintain a constant pressure output to servo valve 34 independent of the diminution of pres sure in source 46. The thrust from jets 36 and 38 is thus controlled as a linear function of the electrical control signal from computer 16 by means of the servo system comprising servo valve 34. manifold block 40. and servo regulators 42.44. It should be noted that jets 36 and 38 are oriented to provide precisely controllable thrust forces normal to the missile longitudinal axis and are located in the same longitudinal plane as the angular position and rate pickoffs l2 and I4. Jets 36 and 38 may be directed outwardly into the airstream or inwardly in the rocket booster nozzle. and can eject gas in the former case. or gas or propellant in the latter. Control can also be maintained by other means well known in the art such as by vanes in the rocket booster exit or by air vanes. In the present embodiment pneumatic jets are utilized. directed out into the airstrcam.

FIG. 2 shows a schematic ofthe control computer 16 which is a semiconductor circuit with the basic function of conditioning the outputs of the pickoffs l2 and [4 for proportional control of servo valve 34. Pickoff I2 is an angular position pickoff whose output is proided as a suppressed carrier modulated signal with the modulation being the angular position error as relleeted in gyro l0. Pickoff I4 is an inductive rate pickoff whose output signal is pro ortional to the rate (degrees per unit time) ofthe angular error as reflected by the gyro [0. These signals form the inputs to computer 16 as shown with d) being the term for position d) the term for rate. The 05' signal is amplified or multiplied by a constant term I) in amplifiers l8 and 22. respectively. Both signals are added in summing amplifier 20 to perform the step a d) h d) which combined output is further amplified in demodulator 24 to provide a control current output to the coil 32 of servo valve 34 which control current is proportional to the error or deviation experienced by the missile in its boost flight front the launch predetermined flight. DC. source 28 furnishes the B and B biases as required for operation of the computer circuitry. The rate signal input 4) is fed as an input to input terminal 49 of a solid state chopper 50 which may be a solid state chopper such as that currently commercially available from the Solid State Electronic Corporation of Van Nuys. California as their Model No. 70. Chopper S0 is driven by a 400 cps reference frequency front reference frequency source 31) through transformer 52 across chopper input terminals 54.56. Chopper 50 operates in the conventional manner to convert the rate input (1) to the chopping frequency of 400 c.p.s. The output of chopper 50 from output terminals 58 and 60 is fed through coupling capacitors 62,64 into amplifier 22 which is a push-pull amplifier including a pair of NPN transistors 66 and 68. Bias for the bases of transistors 66 and 68 is furnished through resistors 70. 72, 77 and common bleeder resistor 74. Emitter bias is furnished by resistors 78. 84 and 82. 86 in the manner shown. Bypass capacitors 75.76 and are utilized to avoid loss in gain. Capacitor 87 is connected across the collectors of transistors 66 and 68 and is used as a peaking capacitor to tune coil 88. The output of amplifier 22 is fed into coupling transformer 88 which provides through its secondary and across resistance 90 and potentiometer 92 the rate signal corresponding to h d).

The output from position pickoff 12 corresponds to the term d). Position pickoff i2 is embodied as a differential transformer whose output voltage is of the order of 1.00 volts per degree of angular movement of gyro l0 and whose output will be in phase with the excitation frequency for one direction oferror input and [80 out of phase for the opposite direction of error input. The position signal is developed across potentiometer 94 and fixed resistor 96. The rate signal through fixed resistor 98 and the position signal through resistor 96 at the primary of transformer 99 and the ratio at which these are summed is determined by the ratio of resistor 98 to resistor 96. Potentiometers 92 and 94 are selectively adjustable as gain control devices. At the summing point at the primary of transformer 99. the rate and position signals which will appear as 400 c.p.s. signals are added. lf the two signals are in phase. they will be added to give a resultant greater than either the rate or position signal as they exist separately. If the two signals are out of phase then only the difference between their relative amplitudes will remain across the primary winding of transformer 99. The composite rate plus position signal is coupled by transformer 99 to a push'pull amplifier stage comprising NPN transistors 100 and 102. Biasing is accomplished by resistors [04. 106.

108. ll(). ll], and I14 as shown. Bypass capacitors H6. H8. I20 are utilized in conjunction with resistors I06, I08. and I I2 as shown to maintain a constant bias. Capacitor I22 is connected across the collectors of transistors I and I02. The output of amplifier which comprises the amplified composite rate position signal is coupled through transformer I24 to a demodulator stage 24 which includes transistors I26 and I28. A reference 400 c.p.s. signal is supplied from reference signal source 30 to the secondary of transformer I24 through a resistor network comprising resistors I30. I32 and potentiometer I34 with the amplitude of the reference signal preset by the setting of potentiometer I34. This reference signal is furnished at the required voltage-frequency for demodulation of the a qb +1) (1) signal. To provide a quiescent state current for the servo valve coil 32 and to provide power for the demodulator transistors I26. I28. The outputs of transistors I26 and I28 are connected through semiconductor diodes I36 and I38, respectively. to servo valve coil 32. The current output from the demodulator stage 24 is a D.(. with a 20 percent 400 cycle per second ripple frequencey due to the inclusion of capacitor I40. Potentiometer 142 is utilized to allow adjustment of the symmetry of current through the two servo coils 32.

FIG. 3 shows the detail of one of the servo regulator valves 42 in a closed position with its associated thrust jet 36 mounted thereon. The valve body I50 has a central hemispheric chamber portion I52 communicating with thrust jet 36 and a second chamber portion I54 formed near its righthand end. A control pressure inlet I56 is connected to the output from servo valve 34 and an operating pressure inlet port I58 is directly connected to the source of pressurized fluid 46. The valve operating member is indicated generally by the numeral I60. Operating member I60 is cylindrical in shape and differentially movable along the axis of valve body 150 in accordance with control pressure from servo valve 34. Located proximate opposite ends of operating member I60 are land portions 162 and I64 comprising driving piston I62 and damping piston I64. respectively, which are seated in central annular chamhers I66 and I54 formed in valve body 150. Annular chamber I66 has a plurality of ports I68 connecting it to central chamber portion 152. Annular chamber I54 has a plurality of ports I70 communicating between operating pressure inlet I58 and chamber I54 or. alternately they may be joined through an enlarged bore through valve body I50. It will further be noted that damping piston I64 has a plurality of orifices I72. one of which is shown. longitudinally formed therein. A vent hole I74 is provided in the right end portion of valve body 150. The actual opening and closing of the valve body is accomplished by a central land portion I76 formed on operating member I60. Rightward movement of operating member I60 serves to pass operating pressure from inlet port I58 to central chamber I52 and out through thrust jet 36.

FIG. 4 shows the manifold block which in connection with seno valve 34 forms the second portion ofthe servo system utilized in the present invention. Bolt holes I78U-t/ are provided for mounting the manifold block 40 to the base of servo valve 34 in the manner illustrated in FIG. 5. hereinafter. Aperture I780 communicates through manifold block 40 to servo valve 34 to provide a pressure vent therefor. The manifold block 40 has an inlet nol/le I80 and a conduit I82 formed therein serving as the input of total pressure from the source of pressurized fluid 46 to the servo valve 34. Output fittings I84 and I86 extend downwardly from the manifold block 40 and are connected through conduits to provide a control pressure for servo regulators 42 and 44, respectively. A pair of orifice fittings I88 and 190 are provided which are connected to vertically extending conduits I92 and I94 which receive the out puts of control pressure from servo valve 34. Orifice fittings I88 and 190 serve to communicate between conduits I92 and I94, respectively. and atmospheric pressure. It is the function ofthe manifold block 40 and servo valve 34 to rapidly generate the control pressure in linear response to an input signal. in the present embodiment as hereinbefore described, an electrical control current. The servo valve 34 utilized may be any of a number of currently. commercially available. electromagnetically operable servo valves such as those produced and sold by Weston Hydraulics Limited. a sub sidiary of Borg-Warner Corporation. Van Nuys. California. The conversion of a flow rate valve to a pressure control device is achieved by providing downstream orifice fittings I88 and I90 which have a size 1.88 times the maximum valve port area provided by conduits I92 and I94. This assures sonic flow into the servo regula tors 42 or 44 during the entire travel of the servo valve operating member 160. Under isothermal flow condi tions the pressure in the servo regulator 42. as shown in FIG. 3, will be directly proportional to flow rate from the servo valve 34 which in turn is proportional to the valve port area as controlled by the electrical signal.

FIGS. 5 through 7 show further detail and the mode of interconnection of conduits I92 and I94 which receive the flow output from servo valve 34. the orifice fitting I88, I90 and the output fittings I84, I86. Also shown is the input fitting I80 in the manifold block 40 which communicates therethrough to provide operat ing pressure for the servo valve 34. The input fitting is coupled directly to the source of pressurized fluid 46 in the manner illustrated in FIG. I.

FIG. 8 shows the detail of position pickoff I2 and its cooperative relationship with the rotor 10a of gyroscope I0. Position pickoft" I2 comprises a core 200 including an excitation winding 202 and a pair of output windings 204,206. An alternating current of 400 cycles per second input is applied across terminals 208. 210 to winding 202. The rotor I011 is shown with its iron ring 212 in juxtaposition to the respective windings on core 200. Depending on whether the pickoff I2 is above or below the reference iron ring 2I2, the coupling between winding 202 and windings 204, 206 will be varied. Accordingly, the amplitude of the voltage induced across the output windings 204, 206, will be varied in a manner proportional to the angle of rotation of the missile from its predetermined attitude. The phase relation of the voltages across the output windings 204, 206 will be indicative of the direction of rotation with reference to gyro rotor I00. The output from leads 2H. 216 is applied across resistor 94 in the man ner hereinbefore indicated in FIG. 2.

FIG. 9 shows the detail of rate pickoff I4 and illustrates the manner in which it is cooperable with the reference iron ring 2I2 of gyro rotor I041 to provide a voltage output which is proportional to the rate of missile rotation from its predetermined attitude. Rate pickoff I4 is similar to pick-off I2 and includes an *E core 300 carrying an excitation winding 302. and output windings 304 and 306 connected as shown. The input to terminals 308. 3I0. however. is a direct current voltage input. A DC flux field results about the center leg of core 300. With the rotor ring 2I2 centered relative to thc pickoff. the flux field remains static and no voltage output occurs across output windings 304. 306. When relative rotation occurs between the pickoff and rotor [a. the DC flux field is cut thereby inducing a voltage across the output windings which is proportional to the rate of rotation of the missile from its predetermined attitude. The voltage across output leads 314. 316 is of one polarity for a motion in one direction and of the opposite polarity for an opposite direction of motion. Leads 314. 316 are connected to terminals 49. 60. respectively. as indicated in FIG. 2.

DESCRIPTION OF OPERATION Upon displacement of the missile from its predetermined attitude. electrical output signals will be generated from the rotor of the gyroscope I0 through position pickoff I2 and rate pickoff I4 as signals (b and (1). respectively.

As shown in FIG. I. it will be seen that the multiplication of the position signal (I) by the constant a and the multiplication of the rate signal (1) by the constant I) together with their summation is accomplished in computer I6. With reference to FIG. 2, the rate signal is modulated to a 400 cps signal by modulator 2] and amplified ie. multiplied by the preset factor b by the push-pull amplifier stage comprising transistors 66. 68. The position signal ql) which is also a 400 c.p.s. signal is summed with h 4) signal across the primary of transformer 99. Amplification of the composite rate position signal is accomplished in the push-pull amplifier stage comprising transistors I00. I02. Demodulation opera tion in the final stage of computer I6 is accomplished in transistors I26, I28 operating as a class B push-pull power amplifier. Transistors I26 and 128 may be medium power silicon transistors such as type 2N497. When its base is rendered positive by the 400 c.p.s. reference signal applied across the secondary of transformer I24, transistor 126 is in a conductive state. A current pulse at 400 c.p.s. is allowed to flow through transistor I26. diode I36 and one-half of servo valve coil 32. It will be noted that transistor 128 has been shut off during this period. On the next cycle of 400 c.p.s. when its base is positive. transistor I28 is turned on and transistor I26 remains off. Operation on conduction of transistor I28 provides a current flow through diode I28 and the other halfof servo valve coil 32. This alternate conduction of transistors I26, I28 results in a pulsating current flow in an opposite direction through each half of the coil. This current may be of the order of approximately (L004 amps in each direction through the coil. Since the current flow is equal and in opposite directions. servo valve 34 will not be actuated and will remain off.

Control signals appearing at the secondary of transformer 124 will be either in or out of phase with the 400 cps reference voltage frequency. The sum of the in phase components will appear across transistor I26 and the difference of the out of phase components will appear across transistor 128. This will result in a net difference between currents through each half of coil 32. Servo valve 34 will thus be cnergiled in accordance with the J or .I signals so that corrective forces pro portional to those signals will be provided through jets 36 and 38. Whether I is plus or minus will determine which of regulators 42 and 44 is activated by control pressure llow front servo valve 34. As best shown in FIG. 3. in the event of a negative signal. control pressure is forwarded from the servo valve 34. through manifold block unit 40, and through conduit I56 to the lefthand thrust face of the driving piston I62 of servo regulator valve 42. This application of control pressure will move the cylinder valve operating member I60 rightwardly. It should be noted that operating or working pressure is at all times furnished through input conduit 158 to the servo regulator valve 42. Prior to the rightward movement of the valve operating member I60, operating pressure has thus been furnished through conduits and apertures I72 in damping piston 164 to provide equal pressure on either thrust face of damping piston 164. As soon as the valve operating member I60 moves rightwardly. land 176 thereon emits operating pressure to the central chamber I52. Operating pressure is then communicated from chamber 152 with a force feedback effect through apertures 16S and into chamber 166 against the righthand thrust face ofdriving piston 162 to exert an opposing force to arrest rightward motion of operating member I60. At the same time. the fluid in righthand chamber 154 is compressed and is forced readily through aperatures 170 and apertures 172 in damping piston I64. The construction of the servo regulator 42 permits a force feedback effect against driving piston I62 and an almost simultaneous damping effect against damping piston 164. It should be further noted that the present mode of construction eliminates the need for mechanical damping devices such as springs which might provide a nonlinearly operating valve. This elimination of mechanical damping elements makes the servo regulator valve 42 a linear pneumatic amplifier whose output is controlled by a constant balance between the input control pressure and the output pressure in chamber I52, and is independent of the magnitude of supply pressure from pressurized fluid source 46. The present invention has additional advantages of utilizing the high density fluid supply as a damping medium. The high density supply gas from source 46 is readily available and its employment requires the use of no special seals in piston I64 since a slight leakage of this fluid causes no problem. The damping chamber 154 formed at the righthand end of valve body ISO is made to act as a degenerative spring. Clearance for operating member 160 is provided by chamber 175 which is vented through port 174.

It will thus be seen that I have provided a pressure regulator valve in which a control pressure from a relatively low-flow source can proportionately control the pressure of a working fluid from a relatively high flow source. The pressure regulator valve is important in the present directional control system in that it operates in a stablized manner independent of changes in fluid supply pressure.

It will thus be seen that by the present invention I have provided a directional control system and method for missiles effective to provide a constant direction of travel during its boost flight. It functions with reference to a simple inertial directional reference in an im proved manner not hitherto possible.

What is claimed is:

I. An attitude control system for maintaining a missile on a launch predetermined path comprising means mounted on the missile for indicating a deviation from said predetermined path including means for providing an output signal representative of angular position change; and means for providing an output signal representative of angular rate change means for modifying said angular position change signal by multiplying it by a factor (1 whereby a E[(TD)/(k/P) ll an algebraic summation means responsive to said modified position and rate change signals and operable to provide a combined output representative of their algebraic summation, and means operatively connected to and controlled by the output ofsaid summation means. said last mentioned means operable to provide a corrective force to return the missile to its predetermined path.

2. An attitude control system for maintaining a mis sile on a launch predetermined path comprising a gyroscope having its spin axis mounted along the longitudi nal axis of the missile at launch. a first and a second pickoff operatively connected to said gyroscope and responsive to movement of the spin axis of said gyroscope from said longitudinal axis, said first pickoff operablc to provide an electrical output representative of position deviation and said second pickoff operable to provide an electrical output representative of rate of deviation. means for modifying the output of said first pickoff by multiplying it by a factor "0 whereby a E|(T-D)/(k/P) I] an algebraic summation means operatively connected to the modified output of said first pickoff and the output of said second pickoff for providing an electrical output signal representative of their algebraic summation, and thrust jet means mounted on said missile and lying in the same control plane as said pickoffs for providing a corrective force proportional to the output signal from said last mentioned means to return the missile to its predetermined path.

3. The combination as set forth in claim 2 in which an electrical servo valve has its control winding operatively connected to the output of said algebraic summation means. and a servo regulator valve is operatively connected between said servo valve and said thrust jet means and a source of constant pressure fluid to provide said corrective force.

4. An attitude control system for maintaining a missile in a launch predetermined path comprising means mounted on the missile for indicating a deviation from said predetermined path including a first pickoff means for providing an electrical output signal representative of angular position change and a second pickoff means for providing an electrical output signal representative of angular rate change, a means operatively connected to each of the aforesaid first and second means for multiplying each of said signals by a constant factor representative ofan average value of gain coefficient as fixed by aerodynamic characteristics and geometry of the missile. means operatively connected to the output of said multiplying means for algebraically summing their outputs. and means operatively connected to and controlled by the output of said summing means. said last mentioned means operable to provide a corrective force to return the missile to its predetermined path.

5. The combination as set forth in claim 4 in which said last-mentioned means comprises a pair of oppositely oriented thrust jet means mounted on the missile in a common control plane with said first and second pickoff means.

6. The combination as set forth in claim 5 in which an electrical servo valve has its control winding operativcly connected to the output of said summing means and a servo regulator is operatively connected to and controlled by said servo valve and connected between said thrust jet means and a source of constant pressure fluid to provide said corrective force.

7. An attitude control system for maintaining a missile on a launch predetermined path comprising a gyroscope having its spin axis fixed along the longitudinal axis of the missile at launch, a first and a second pickoff operatively connected to said gyroscope and responsive to relative movement between the spin axis of said gyroscope and said longitudinal axis, said first pickoff operable to provide an electrical output representative of position deviation and said second pickoff operable to provide an electrical output representative of rate of deviation, means for multiplying the output of each of said pickoffs by a different constant factor representative of an average value of gain coefficient as fixed by aerodynamic characteristics and geometry of the mis sile. an algebraic summation means operatively connected to the outputs of said multiplying means for pro viding an electrical output signal representative of their summation, means for applying a corrective force normal to the longitudinal axis of said missile, said last mentioned means operable responsive to the output of said summation means to provide the corrective force in a direction normal to said longitudinal axis of said missile.

8. The combination as set forth in claim 7 in which said means for applying a corrective force comprises a pair of oppositely oriented thrust jets mounted on said missile, both of said jets lying in a common control plane with said first and second pickoff means.

9. The combination as set forth in claim 8 in which an electrical servo valve has its control winding operatively connected to the output of said summation means, and a servo regulator valve is operatively connected between said servo valve and a source of constant pressure fluid to provide said corrective force through said jets.

[0. The method of maintaining a missile on a launch predetermined path during boost flight comprising. de riving a signal representative of the missiles deviation from said predetermined path. deriving a signal representative of the rate of the missiles deviation from said path. multiplying each of said signals by its own constant factor representative of an average value of gain coefficient as fixed by aerodynamic characteristics and geometry of the missile, algebraically summing both the resultant signals. and utilizing the composite signal to apply and control the magnitude of a corrective force to the missile.

11. The method of maintaining a missile on a launch predetermined path during boost flight comprising. deriving from an inertial reference a signal representative of the missiles deviation from said path. deriving from said inertial reference a signal representative of the rate ofthe deviation from said path. multiplying each of said signals by its own constant factor representative of an average value of gain coefficient as fixed by the aerodynamic characteristics and geometry of the missile. algebraically summing both the resultant signals. and applying a corrective jet force normal to the longitudinal axis of the missile proportional to the summed signal.

12. The method of maintaining a missile on a launch predetermined path during boost flight comprising. dcriving from an inertial reference a signal representative ofthe missile's deviation from said predetermined path. deriving from said inertial reference a signal representative of the rate of the missiles deviation from said path. multiplying each of said signals by its own constant factor representative of an average value of gain coefficient as fixed by the aerodynamic characteristics and geometry of the missile algebraically summing both the resultant signals, and utilizing the composite signal to apply and control the magnitude of a correc tive force to the missile.

13. The method of maintaining a missile on a launch predetermined path during boost flight comprising. deriving from an inertial reference an electrical signal representative of the missilc's deviation from said predetermined path. deriving from said inertial reference an electrical signal representative ofthe rate of the missiles deviation from said path. modulating each of said signals to a common reference frequency. multiplying each of said signals by its own constant factor representative of an average value of gain coefficient as fixed by the aerodynamic characteristics and geometry of the missile, algebraically summing both of the resultant signals, demodulating the composite signal and utilizing the composite signal to apply and control the magnitude of a corrective force to the missile.

14. The method of maintaining a missile on a launch predetermined path during boost flight comprising. deriving a signal representative of the missile deviation from said predetermined path. deriving a signal representative of the rate of the missile deviation from said path multiplying at least one of the aforesaid signals by (all a constant factor representative of an average value of gain coefficient as fixed by aerodynamic characteristics and geometry of the missile. algebraically summing the resultant signals and utilizing the composite signal to apply and control the magnitude of a corrective force to the missile.

15. The method of maintaining a missile on a launch predetermined path during boost flight comprising. deriving a signal representative of the missile deviation from said predetermined path. deriving a signal representative of the rate of the missile deviation from said path. multiplying at least one of the aforesaid signals by a constant factor 0" whereby a E[(TDil(k/Pl l algebraically summing the resultant signals. and utilizing the composite signal to apply and control the magnitude of a corrective force to thee missile.

16. An attitude control system for maintaining a missile in a launch predetermined path comprising means mounted on the missile for indicating its deviation from said predetermined path including a first pickoff means for providing an electrical output signal representative of angular position change and a second pickoff means for providing an electrical output signal rcpresentathe of angular rate change a means operatively connected to the output signal of one of said pickoff means for multiplying it by a constant factor "11" whereby a -=-[(T-D)/(k/p)- ll. a summing means operatively connected to the output signal of said multiplying means and the output signal of the other of said pickoff means. and means operatively connected to and controlled by the output of said summing means for providing a corrective force to return the missile to its predetermined path. 

1. An attitude control system for maintaining a missile on a launch predetermined path comprising means mounted on the missile for indicating a deviation from said predetermined path including means for providing an output signal representative of angular position change; and means for providing an output signal representative of angular rate change means for modifying said angular position change signal by multiplying it by a factor ''''a'''' whereby a congruent ((T-D)/(k/p) - 1 ), an algebraic summation means responsive to said modified position and rate change signals and operable to provide a combined output representative of their algebraic summation, and means operatively connected to and controlled by the output of said summation means, said last mentioned means operable to provide a corrective force to return the missile to its predetermined path.
 2. An attitude control system for maintaining a missile on a launch predetermined path comprising a gyroscope having its spin axis mounted along the longitudinal axis of the missile at launch, a first and a second pickoff operatively connected to said gyroscope and responsive to movement of the spin axis of said gyroscope from said longitudinal axis, said first pickoff operable to provide an electrical output representative of position deviation and said second pickoff operable to provide an electrical output representative of rate of deviation, means for modifying the output of said first pickoff by multiplying it by a factor ''''a'''' whereby a Congruent ((T-D)/(k/p) - 1 ), an algebraic summation means operatively connected to the modified output of said first pickoff and the output of said second pickoff for providing an electrical output signal representative of their algebraic summation, and thrust jet means mounted on said missile and lying in the same control plane as said pickoffs for providing a corrective force proportional to the output signal from said last mentioned means to return the missile to its predetermined path.
 3. The combination as set forth in claim 2 in which an electrical servo valve has its control winding operatively connected to the output of said algebraic summation means, and a servo regulator valve is operatively connected between said servo valve and said thrust jet means and a source of constant pressure fluid to provide said corrective force.
 4. An attitude control system for maintaining a missile in a launch predetermined path comprising means mounted on the missile for indicating a deviation from said predetermiNed path including a first pickoff means for providing an electrical output signal representative of angular position change and a second pickoff means for providing an electrical output signal representative of angular rate change, a means operatively connected to each of the aforesaid first and second means for multiplying each of said signals by a constant factor representative of an average value of gain coefficient as fixed by aerodynamic characteristics and geometry of the missile, means operatively connected to the output of said multiplying means for algebraically summing their outputs, and means operatively connected to and controlled by the output of said summing means, said last mentioned means operable to provide a corrective force to return the missile to its predetermined path.
 5. The combination as set forth in claim 4 in which said last-mentioned means comprises a pair of oppositely oriented thrust jet means mounted on the missile in a common control plane with said first and second pickoff means.
 6. The combination as set forth in claim 5 in which an electrical servo valve has its control winding operatively connected to the output of said summing means and a servo regulator is operatively connected to and controlled by said servo valve and connected between said thrust jet means and a source of constant pressure fluid to provide said corrective force.
 7. An attitude control system for maintaining a missile on a launch predetermined path comprising a gyroscope having its spin axis fixed along the longitudinal axis of the missile at launch, a first and a second pickoff operatively connected to said gyroscope and responsive to relative movement between the spin axis of said gyroscope and said longitudinal axis, said first pickoff operable to provide an electrical output representative of position deviation and said second pickoff operable to provide an electrical output representative of rate of deviation, means for multiplying the output of each of said pickoffs by a different constant factor representative of an average value of gain coefficient as fixed by aerodynamic characteristics and geometry of the missile, an algebraic summation means operatively connected to the outputs of said multiplying means for providing an electrical output signal representative of their summation, means for applying a corrective force normal to the longitudinal axis of said missile, said last-mentioned means operable responsive to the output of said summation means to provide the corrective force in a direction normal to said longitudinal axis of said missile.
 8. The combination as set forth in claim 7 in which said means for applying a corrective force comprises a pair of oppositely oriented thrust jets mounted on said missile, both of said jets lying in a common control plane with said first and second pickoff means.
 9. The combination as set forth in claim 8 in which an electrical servo valve has its control winding operatively connected to the output of said summation means, and a servo regulator valve is operatively connected between said servo valve and a source of constant pressure fluid to provide said corrective force through said jets.
 10. The method of maintaining a missile on a launch predetermined path during boost flight comprising, deriving a signal representative of the missile''s deviation from said predetermined path, deriving a signal representative of the rate of the missile''s deviation from said path, multiplying each of said signals by its own constant factor representative of an average value of gain coefficient as fixed by aerodynamic characteristics and geometry of the missile, algebraically summing both the resultant signals, and utilizing the composite signal to apply and control the magnitude of a corrective force to the missile.
 11. The method of maintaining a missile on a launch predetermined path during boost flight comprising, deriving from an inertial reference a signal representative of the missile''s deviation From said path, deriving from said inertial reference a signal representative of the rate of the deviation from said path, multiplying each of said signals by its own constant factor representative of an average value of gain coefficient as fixed by the aerodynamic characteristics and geometry of the missile, algebraically summing both the resultant signals, and applying a corrective jet force normal to the longitudinal axis of the missile proportional to the summed signal.
 12. The method of maintaining a missile on a launch predetermined path during boost flight comprising, deriving from an inertial reference a signal representative of the missile''s deviation from said predetermined path, deriving from said inertial reference a signal representative of the rate of the missile''s deviation from said path, multiplying each of said signals by its own constant factor representative of an average value of gain coefficient as fixed by the aerodynamic characteristics and geometry of the missile, algebraically summing both the resultant signals, and utilizing the composite signal to apply and control the magnitude of a corrective force to the missile.
 13. The method of maintaining a missile on a launch predetermined path during boost flight comprising, deriving from an inertial reference an electrical signal representative of the missile''s deviation from said predetermined path, deriving from said inertial reference an electrical signal representative of the rate of the missile''s deviation from said path, modulating each of said signals to a common reference frequency, multiplying each of said signals by its own constant factor representative of an average value of gain coefficient as fixed by the aerodynamic characteristics and geometry of the missile, algebraically summing both of the resultant signals, demodulating the composite signal, and utilizing the composite signal to apply and control the magnitude of a corrective force to the missile.
 14. The method of maintaining a missile on a launch predetermined path during boost flight comprising, deriving a signal representative of the missile deviation from said predetermined path, deriving a signal representative of the rate of the missile deviation from said path, multiplying at least one of the aforesaid signals by a constant factor representative of an average value of gain coefficient as fixed by aerodynamic characteristics and geometry of the missile, algebraically summing the resultant signals and utilizing the composite signal to apply and control the magnitude of a corrective force to the missile.
 15. The method of maintaining a missile on a launch predetermined path during boost flight comprising, deriving a signal representative of the missile deviation from said predetermined path, deriving a signal representative of the rate of the missile deviation from said path, multiplying at least one of the aforesaid signals by a constant factor '''' whereby a Congruent ((T-D)/(k/p) - 1 ), algebraically summing the resultant signals, and utilizing the composite signal to apply and control the magnitude of a corrective force to thee missile.
 16. An attitude control system for maintaining a missile in a launch predetermined path comprising means mounted on the missile for indicating its deviation from said predetermined path including a first pickoff means for providing an electrical output signal representative of angular position change and a second pickoff means for providing an electrical output signal representative of angular rate change, a means operatively connected to the output signal of one of said pickoff means for multiplying it by a constant factor ''''a'''' whereby a Congruent ((T-D)/(k/p) - 1 ), a summing means operatively connected to the output signal of said multiplying means and the output signal of the other of said pickoff means, and means operatively connected to and controlled by the output of said summing means for providing a corrective force to Return the missile to its predetermined path. 